Vane arc segment with thermal insulation element

ABSTRACT

Disclosed is a method of reducing play in a vane arc segment. The vane arc segment includes an airfoil piece that defines first and second platforms and a hollow airfoil section that has an internal cavity and that extends between the first and second platforms. The first platform defines a gaspath side, a non-gaspath side, and a radial flange that projects from the non-gaspath side. Support hardware supports the airfoil piece via the radial flange, and a thermal insulation element is located adjacent the radial flange. The method includes performing a light scan of the radial flange to produce a digital three-dimensional model of the radial flange, and then machining the thermal insulation element in accordance with the digital three-dimensional model to provide a low-tolerance fit between the radial flange and the thermal insulation element that limits play between the airfoil piece and the thermal insulation element.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section may include low and high pressure compressors, and the turbine section may also include low and high pressure turbines.

Airfoils in the turbine section are typically formed of a superalloy and may include thermal barrier coatings to extend temperature capability and lifetime. Ceramic matrix composite (“CMC”) materials are also being considered for airfoils. Among other attractive properties, CMCs have high temperature resistance. Despite this attribute, however, there are unique challenges to implementing CMCs in airfoils.

SUMMARY

A vane arc segment according to an example of the present disclosure includes an airfoil piece that defines first and second platforms and a hollow airfoil section that has an internal cavity and that extends between the first and second platforms. The first platform defines a gaspath side, a non-gaspath side, and a radial flange that projects from the non-gaspath side. Support hardware supports the airfoil piece via the radial flange, and a thermal insulation element is located adjacent the radial flange. A method of reducing play in the vane arc segment includes performing a light scan of the radial flange to produce a digital three-dimensional model of the radial flange, and then machining the thermal insulation element in accordance with the digital three-dimensional model to provide a low-tolerance fit between the radial flange and the thermal insulation element that limits play between the airfoil piece and the thermal insulation element.

In a further embodiment of the foregoing embodiment the airfoil piece and the thermal insulation element are ceramic.

In a further embodiment of any of the foregoing embodiments, the light scan is a structured light scan.

In a further embodiment of any of the foregoing embodiments, the thermal insulation element includes a coating, and the machining of the thermal insulation element includes machining the coating.

In a further embodiment of any of the foregoing embodiments, the thermal insulation element is a monolithic ceramic or a ceramic matrix composite.

In a further embodiment of any of the foregoing embodiments, the thermal insulation element includes a silicon-containing ceramic.

In a further embodiment of any of the foregoing embodiments, the thermal insulation element includes a coating.

In a further embodiment of any of the foregoing embodiments, the coating includes at least one of hafnia, silica, silicate, or zirconia.

In a further embodiment of any of the foregoing embodiments, the machining includes machining of the coating.

In a further embodiment of any of the foregoing embodiments, the performing of the light scan includes a structured light scan by blue light.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the present disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows.

FIG. 1 illustrates a gas turbine engine.

FIG. 2 illustrates a portion of a vane arc segment.

FIG. 3 illustrates an isolated view of an airfoil piece.

FIG. 4 illustrates a portion of another example vane arc segment.

FIG. 5 illustrates another example vane arc segment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. The fan section 22 drives air along a bypass flow path B in a bypass duct defined within a housing 15 such as a fan case or nacelle, and also drives air along a core flow path C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 that interconnects, a first (or low) pressure compressor 44 and a first (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive a fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a second (or high) pressure compressor 52 and a second (or high) pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 may be arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded through the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of the low pressure compressor, or aft of the combustor section 26 or even aft of turbine section 28, and fan 42 may be positioned forward or aft of the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1 and less than about 5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 meters/second).

FIG. 2 illustrates a sectioned view (taken mid-chord, aft looking forward) through a portion of a vane arc segment 60 of a vane ring assembly from the turbine section 28 of the engine 20. The vane arc segments 60 are situated in a circumferential row about the engine central axis A. Although the vane arc segment 60 is shown and described with reference to application in the turbine section 28, it is to be understood that the examples herein are also applicable to structural vanes in other sections of the engine 20.

The vane arc segment 60 is comprised of an airfoil piece 62, which is also shown in a full isolated view in FIG. 3 . The airfoil piece 62 includes several sections, including first and second platforms 64/66 and an airfoil section 68 that extends between the first and second platforms 64/66. The airfoil section 68 defines a leading edge 68 a, a trailing edge 68 b, and pressure and suction sides 68 c/68 d. The airfoil section 68 generally circumscribes an internal cavity 70 such that the airfoil section 68 in this example is hollow. The terminology “first” and “second” as used herein is to differentiate that there are two architecturally distinct components or features. It is to be further understood that the terms “first” and “second” are interchangeable in the embodiments herein in that a first component or feature could alternatively be termed as the second component or feature, and vice versa.

In this example, the first platform 64 is a radially outer platform and the second platform 66 is a radially inner platform relative to the engine central longitudinal axis A. The first platform 64 defines a gaspath side 64 a and a non-gaspath side 64 b. Likewise, the second platform 66 defines a gaspath side 66 a and a non-gaspath side 66 b. The gaspath sides 64 a/66 a bound the core flow path C through the engine 20.

The platform 64 further includes a radial flange 72 that projects (radially) from the non-gaspath side 64 b. In this example, the flange 72 is an airfoil-shaped collar that is in essence an extension of the airfoil section 68 radially past the platform 64. It is to be appreciated, however, that the examples herein may be applied to flanges having other geometries that are also the exclusive load-transmitting attachments of an airfoil piece. The flange 72 defines an inner face 72 a that generally faces toward the internal cavity 70, an opposed outer face 72 b, and a radial face 72 c. The flange 72 serves to transfer loads, such as aerodynamic forces, from the airfoil piece 62 to support hardware 74. Likewise, the platform 66 may also include a flange (not shown) that engages support hardware. The airfoil piece 62 is supported exclusively through the flanges 72, i.e., the flanges are the exclusive load-transmitting attachments of the airfoil piece 62.

The airfoil piece 62 is continuous in that the platforms 64/66 and airfoil section 68 constitute a one-piece body. As an example, the airfoil piece 62 is formed of a ceramic material, an organic matrix composite (OMC), or a metal matrix composite (MMC). For instance, the ceramic material is a monolithic ceramic or a ceramic matrix composite (CMC) that is formed of ceramic fibers that are disposed in a ceramic matrix. The monolithic ceramic may be, but is not limited to, SiC or other silicon-containing ceramic. The ceramic matrix composite may be, but is not limited to, a SiC/SiC ceramic matrix composite in which SiC fibers are disposed within a SiC matrix. Example organic matrix composites include, but are not limited to, glass fiber, carbon fiber, and/or aramid fibers disposed in a polymer matrix, such as epoxy. Example metal matrix composites include, but are not limited to, boron carbide fibers and/or alumina fibers disposed in a metal matrix, such as aluminum. The fibers may be provided in fiber plies, which may be woven or unidirectional and may collectively include plies of different fiber weave configurations.

The airfoil piece 62 is mounted between support hardware 74 at both the inner and outer diameter ends of the airfoil piece. In the example of FIG. 2 , the support hardware 74 is a spar that includes a spar platform 74 a and a leg 74 b. The leg 74 b extends radially from the spar platform 74 a through the internal cavity 70 of the airfoil section 68. The leg 74 b may extend radially past the second platform 66, where it is secured with the support hardware at the inner diameter. In this example, the leg 74 b is hollow and may be provided with pass-through air for cooling downstream components and/or cooling air used to cool a portion of the airfoil piece 62. The support hardware 74 is formed of a metallic alloy that can bear the loads received, such as nickel- or cobalt-based superalloys.

In general, the materials contemplated for the airfoil piece 62 have significantly lower thermal conductivity than superalloys and do not possess the same strength and ductility characteristics, making them more susceptible to distress from thermal gradients and the thermally induced stresses those cause. The high strength and toughness of superalloys permits resistance to thermal stresses, whereas by comparison materials such as ceramics are more prone to distress from thermal stress. Thermal stresses may cause distress at relatively weak locations, such as interlaminar interfaces between fiber plies where there are no fibers carrying load. Therefore, although maximized cooling may be desirable for superalloy vanes, cooling in some locations for non-superalloy vanes may exacerbate thermal gradients and thus be counter-productive to meeting durability goals.

In particular in the vane arc segment 60, the metallic support hardware 74 is cooled by bleed air from the compressor section 24. On the other hand, the airfoil piece 62, most of which is exposed in the core gaspath of the engine 20, operates at relatively high temperatures. As a result, there is the tendency for the metallic support hardware 74 to remove heat from the airfoil piece 62, and thereby increasing thermal gradients across the flange 72 and platform 64. For example, if the flange 72 of the airfoil piece 62 were mounted to be in contact with the support hardware 74, the contact interface would serve as a thermal conductance path for heat removal. For such contact, the flange 72 would be relatively cool in the region of contact. However, since the material from which the airfoil piece 62 is formed has relatively low thermal conductivity, the flange 72 may remain relatively hot in regions that are close the region of contact.

In this regard, as shown in FIG. 2 , the vane arc segment 60 includes a thermal insulation element 76 adjacent the flange 72. The thermal insulation element 76 is located between the flange 72 and the support hardware 74 such that the support hardware 74 supports the airfoil piece 62 through the thermal insulation element 76. The thermal insulation element 76 thus mechanically intervenes between the flange 72 and the support hardware 74 such that there is no direct thermal conductance path via contact. Moreover, the thermal insulation element 76 itself is formed of a low thermal conductance material and thus also limits indirect conductance from the airfoil piece 62 to the support hardware 74. In this regard, the thermal insulation element 76 serves as a mechanical load block and as a thermal insulator. For example, the thermal insulation element 76 is selected from the materials discussed above for the airfoil piece 62. In the turbine section 28, however, the thermal insulation element 76 is ceramic, such as the monolithic ceramic or CMC discussed above.

Referring also to FIG. 3 , the thermal insulation element 76 circumscribes the flange 72. In this example, the thermal insulation element 76 is a two-piece split ring. It is to be understood, however, that the thermal insulation element 76 could alternatively be provided in more pieces, or as a single piece. The thermal insulation element 76 defines an inner face 76 a that generally faces toward the internal cavity 70 of the airfoil piece 62 and an opposed outer face 76 b. The inner face 76 a serves as a bearing face that is in contact with the outer face 72 b of the flange 72. The outer face 76 b also serves as a bearing face and is in contact with the spar platform 74 a. The thermal insulation element 76 is thus trapped between the spar platform 74 a and the flange 72. For example, the thermal insulation element 76 is otherwise unsecured, via bonding or the like, to either the airfoil piece 62 or the support hardware 74.

Optionally, the thermal insulation element 76 includes a coating 78 on its inner face 76 a. The coating 78 serves to provide a relatively smooth, low-wear contact surface for engagement with the flange 72. The coating 78 may also serve for additional thermal insulation. As examples, the coating 78 may be formed of or include hafnia, silica, silicate, or zirconia.

The example in FIG. 2 is at the outer diameter end of the vane arc segment 60. It is to be understood that it may also be applied to the inner diameter end. Likewise, FIG. 4 shows an inner diameter end of a portion of another example vane arc segment 160 but may alternatively be applied to the outer diameter end. In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding elements.

In the example in FIG. 4 , relative to a radial direction (R) of the airfoil piece 62, the flange 172 is sloped such that its inner and outer faces 172 a/172 b are sloped. The inner face 176 a of the thermal insulation element 176 is also sloped such that the faces 172 b/176 a meet at a sloped interface. For instance, both faces 172 a/176 a are sloped at an angle (SA) of 10° to 50°. Additionally, the there is a gap G between the tip of the flange 172 and the support hardware 174 such that the airfoil piece 62 does not bottom out in contact with the support hardware 174. Thus, all of the load transmitted from the airfoil piece 62 to the support hardware 74 is transmitted through the sloped interface. The effect of such transmission through the sloped interface is that the thermal insulation element 176 receives aerodynamic radial, tangential and axial loads, which places the thermal insulation element 176 in compression relative to the engine radial, axial and tangential directions.

FIG. 5 shows a further example vane arc segment 260 that is a combination of the examples of FIGS. 2 and 4 , whereby the above discussions of each are hereby incorporated into the present example. In this example, relative to the engine central axis A, the support hardware 74 is outer diameter support hardware, the support hardware 174 is inner diameter support hardware, and the leg 74 b is secured to the support hardware 174, such as by fastener 80. The fastener 80 is not particularly limited and may be, but is not limited to, a lock pin.

The thermal insulation element 76/176 also provides a method for reducing play in the vane arc segment 60/160/260. In particular, the position of the airfoil piece 62 relative to the fixed support hardware 74/174 affects the flow area between adjacent airfoil sections 68 and aerodynamic loading on the vane arc segments 60. For instance, if there is play (free motion) between the airfoil piece 62 and the support hardware 74/174, aerodynamic loading may cause the airfoil piece 62 to rotate slightly, which could alter the flow area and aerodynamic loads. Thus, reducing play facilitates reductions in variations of the flow area and aerodynamic loading.

An example method of reducing play includes performing a light scan of the radial flange 72/172 to produce a digital three-dimensional model of the radial flange 72/172. For example, the light scan involves a structured light scan, such as by blue light. The digital three-dimensional model of the radial flange 72/172 may be used to verify that the flange 72/172 meets dimensional tolerances. The digital three-dimensional model, however, may also be used in the fabrication of the thermal insulation element 76/176. In this regard, the thermal insulation element 76/176 is machined in accordance with the digital three-dimensional model to provide a low-tolerance fit between the radial flange 72/172 and the thermal insulation element 76/176 that limits play.

As an example, a surface of one component has subtle variations in contour, while a surface of a mating component does not typically have the same contours. As a result, there are variations of how the surfaces mate from part-to-part, and the components must be made with tolerances to account for such variations and ensure fit.

The surface contours of the flange 72/172, however, are represented in the digital three-dimensional model. Thus, the mating surface of the thermal insulation element 76/176 is machined to match the contours of the flange 72/172. Since the contours of the mating surfaces match, there is a closer fit. With such a closer fit, tolerances can be reduced, which ultimately reduces play. It is to be understood that for purposes herein the coating 78, if used, is considered to be a part of the thermal insulation element 76/176. Thus, the inner face 76 a/176 a of the thermal insulation element 76/176 may be machined or, if the coating 78 is used, the coating 78 may be machined. That is, the machining of the thermal insulation element 76/176 can be done after insulating coating 78 is applied.

Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments.

The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims. 

What is claimed is:
 1. A method of reducing play in a vane arc segment, the vane arc segment including an airfoil piece that defines first and second platforms and a hollow airfoil section that has an internal cavity and that extends between the first and second platforms, the first platform defines a gaspath side, a non-gaspath side, and a radial flange that projects from the non-gaspath side, support hardware that supports the airfoil piece via the radial flange, and a thermal insulation element adjacent the radial flange, method comprising: performing a light scan of the radial flange to produce a digital three-dimensional model of the radial flange; and machining the thermal insulation element in accordance with the digital three-dimensional model to provide a low-tolerance fit between the radial flange and the thermal insulation element that limits play between the airfoil piece and the thermal insulation element.
 2. The method as recited in claim 1, wherein the airfoil piece and the thermal insulation element are ceramic.
 3. The method as recited in claim 1, wherein the light scan is a structured light scan.
 4. The method as recited in claim 1, wherein the thermal insulation element includes a coating, and the machining of the thermal insulation element includes machining the coating.
 5. The method as recited in claim 1, wherein the thermal insulation element is a monolithic ceramic or a ceramic matrix composite.
 6. The method as recited in claim 5, wherein the thermal insulation element includes a silicon-containing ceramic.
 7. The method as recited in claim 6, wherein the thermal insulation element includes a coating.
 8. The method as recited in claim 7, wherein the coating includes at least one of hafnia, silica, silicate, or zirconia.
 9. The method as recited in claim 7, wherein the machining includes machining of the coating.
 10. The method as recited in claim 1, wherein the performing of the light scan includes a structured light scan by blue light. 